CN111563292B - Laminar flow airfoil type Re number effect correction method based on flow transition - Google Patents

Abstract

The invention relates to the technical field of airplane pneumatic data acquisition, and discloses a laminar flow airfoil type Re number effect correction method based on flow transition, which mainly comprises two steps of wing characteristic analysis and data correction; specifically, the method comprises the steps of firstly, acquiring the aerodynamic characteristics of the laminar flow wing aircraft within the full Re number range by using a CFD simulation technology, determining the critical Re of the laminar flow wing aircraft, and analyzing the wing characteristics; then, acquiring the Re number effect correction quantity of the laminar flow wing type airplane step by step through a means of combining a pressurization test and a forced transition test; and finally, finishing the Re number effect correction of the wind tunnel test data according to the Re number effect correction quantity. The Re number effect correction method provided by the invention considers the influence of the Re number on the wing flow transition position, further influences the aircraft lift force, resistance and longitudinal moment, effectively improves the accuracy of the Re number effect correction of the laminar flow wing aircraft, and ensures the flight safety.

Description

Laminar flow airfoil type Re number effect correction method based on flow transition

Technical Field

The invention relates to the technical field of airplane pneumatic data acquisition, in particular to a laminar flow airfoil type Re number effect correction method based on flow transition.

Background

Wind tunnel tests are the most direct and effective means for acquiring aerodynamic data of an airplane. The wind tunnel test generally uses a scaling model, the size of the wind tunnel model is reduced, the wind tunnel flow field is different from the inherent characteristics of the real atmosphere, and the like, so that the Re number of the wind tunnel test is greatly different from the real flying Re number of the airplane. Therefore, the Re number effect correction must be carried out on the wind tunnel test data.

At present, the domestic has a relatively mature Re number effect correction method for manned fighters. The method is a Re number effect correction method of manned fighters summarized in the 70 th century in the China aerodynamic research and development center on the basis of combining the research results of domestic airplane flight tests and wind tunnel tests by referring to foreign experiences. The fighter mainly adopts thin wing profiles, and the boundary layer on the surface of the wing is similar to full turbulence when flying in the atmosphere. The Re number mainly has great influence on the surface friction resistance of the wing, and basically does not influence the surface flow state of the wing, so that the influence on the lift characteristic, the longitudinal moment characteristic and the like of the airplane can be ignored. In the test data correction, the Re number correction is generally performed only for the minimum resistance. Generally, during a wind tunnel test, approximate full turbulence of a model is realized by manually pasting a transition tape on the front edge of a wing, so that the state of a boundary layer of the model is consistent with that of a real aircraft, the influence of the Re number on the minimum resistance of the whole aircraft is simplified into the influence of the Re number on the non-compressible resistance of the whole aircraft under the full turbulence state, the Re number and the surface friction resistance are in a double-logarithmic curve relationship, and the model can be obtained by extrapolation through a pressurization test means.

Nowadays, civil aircrafts, high-altitude and high-speed unmanned planes and the like generally adopt laminar flow wing sections with high lift-drag ratio or supercritical laminar flow composite wing sections in order to pursue high cruising efficiency. The laminar flow airfoil profile has the appearance characteristics that the radius of a front edge is small, the upper surface is flat, and the position with the maximum thickness is more back, so that a proper pressure gradient can be established on the airfoil surface, the unstable factors in a laminar flow boundary layer are effectively inhibited, the transition is delayed, and the laminar flow with 40-70% chord length can be generally maintained on the surface of the laminar flow airfoil profile at a design point (cruise state). Research shows that the drag of the laminar wing can be reduced by more than half compared with the drag of the common turbulent wing at the design point. However, when the deviation from the design point is large, the natural laminar flow region tends to disappear, and the resistance rapidly increases. Experimental research shows that the Re effect of the laminar flow airfoil presents complex non-monotonicity, the Re number influences the flow field structure and the pressure distribution of the laminar flow airfoil, and further influences the macroscopic aerodynamic characteristics of the airplane, and in addition, the Re number can also influence the transition position of the laminar flow airfoil, and further influences the lift force, the resistance and the longitudinal moment of the airplane. Therefore, the method for correcting the Re number effect of the fighter plane is not suitable for the plane adopting the laminar flow wing profile.

Disclosure of Invention

The invention provides a laminar flow wing section Re number effect correction method based on flow transition in order to make up the technical blank that the existing Re number effect correction method cannot meet the Re number effect correction of a laminar flow wing section unmanned aerial vehicle, and compared with the Re number effect correction method of the traditional manned fighter plane without considering the influence of wing surface flow state, the Re number effect correction method provided by the invention considers the influence of Re number on the flow transition position of the wing, further influences the lift force, the resistance and the longitudinal moment of the plane, effectively improves the accuracy of the Re number effect correction of the laminar flow wing section plane and ensures the flight safety.

The invention is realized by the following technical scheme:

a laminar flow airfoil type Re number effect correction method based on flow transition mainly comprises two steps of wing characteristic analysis and data correction; specifically, the method comprises the steps of firstly, acquiring the aerodynamic characteristics of the laminar flow wing aircraft within the full Re number range by using a CFD simulation technology, determining the critical Re of the laminar flow wing aircraft, and analyzing the wing characteristics; then, acquiring the Re number effect correction quantity of the laminar flow wing type airplane step by step through a means of combining a pressurization test and a forced transition test; and finally, finishing the Re number effect correction of the wind tunnel test data according to the Re number effect correction quantity.

A laminar flow airfoil type Re number effect correction method based on flow transition comprises the following steps:

step S100: analyzing the characteristics of the wing;

step S200: and (6) data correction.

The wing characteristic analysis of step S100 specifically includes the steps of:

step S110: determining the flight Re number range of the airplane according to the flight envelope of the airplane;

step S120: calculating aerodynamic characteristics of laminar flow wings within the flight Re number range by using a CFD simulation technology to obtain the flow area Xn.jy of the upper surface layer of the wings in a real flight state_{(cruise. alpha.)}Variation curve along with Re number, wing lift coefficient C_{L (cruise alpha)}Variation curve along with Re number and wing resistance coefficient C_{D (cruise alpha)}Variation curve along with Re number and wing lift-drag ratio K_{(cruise. alpha.)}The variation curve with the number of Re;

step S130: observing each change curve in the step S120, and selecting a critical Re number from a plurality of Re numbers;

wherein the critical Re number simultaneously satisfies the following three conditions:

condition a: when the number of the laminar flow areas is less than the critical Re number, the laminar flow area of the upper surface of the wing basically keeps unchanged, but the aerodynamic performance of the wing is poor;

condition b: along with the increase of the Re number, diffusion separation and diffusion-shock wave separation in the airflow boundary layer are gradually eliminated, the lift force is increased, the resistance is reduced, and the lift-drag ratio is improved;

condition c: after the number of the Re is larger than the critical Re number, the laminar flow area is reduced along with the increase of the Re number, the laminar flow transition point gradually moves forwards, the lift force is reduced, the resistance is increased, the lift-drag ratio is reduced, and finally the upper surface of the wing is changed into a full turbulent flow state.

In the data correction process of step S200, the flight envelope of the aircraft is first determined according to the real flight height and the real flight Re number corresponding to the flight M number, and then one of the following manners is selected according to the magnitude relationship between the real flight Re number and the critical Re number to obtain the Re number effect correction amount of the laminar flow airfoil aircraft:

the first method is as follows: if the real flight Re number is smaller than the critical Re number, the test Re number is increased to the real flight Re number or is closer to the real flight Re number by adopting a pressurization test during the test, and the laminar flow wing type airplane Re number effect correction amount in the test data is directly obtained;

the second method comprises the following steps: if the real flying Re number is larger than or equal to the critical Re number, increasing the test Re number to the critical Re number or closer to the critical Re by adopting a pressurization test, and taking the data as reference data; and performing a forced transition test by manually pasting transition belts at different chord-wise positions on the surface of the wing to obtain lift force, resistance and longitudinal torque increment caused by transition position change, and superposing the lift force, the resistance and the longitudinal torque increment on reference data to obtain Re number effect correction of the laminar flow wing type airplane in the test data.

Further, the Re number effect correction comprises a lift coefficient C_LModified, minimum drag coefficient C_DminCorrection, longitudinal moment coefficient C_mAnd (6) correcting.

First, the lift coefficient C_LThe specific steps of the correction are as follows:

step A1: increasing proper pressure in the test process to increase the Re number of the test to the critical Re number or more to the critical Re number, and acquiring the lift coefficient C of the reference state_{L _ boost}；

Step A2: determining the corrected flight height and flight M number, calculating the corresponding real Re number, and obtaining the free transition position Xn corresponding to the real Re number by combining the change curve of the laminar flow region of the upper surface of the wing along with the Re number in the real flight state_{Free transition (CFD)}Simultaneously obtaining the free transition position Xn corresponding to the critical Re number_{Free transition critical}To find the free transition position Xn_{Free transition (CFD)}And a free transition position Xn_{Free transition critical}The difference Δ Xn of (d);

step A3: carrying out normal-pressure wind tunnel test on the laminar flow wing type airplane to obtain the aerodynamic force characteristic of the airplane in the free transition state and obtain the corresponding lift line slope C_{L alpha free transition normal pressure test}And zero lift angle of attack alpha_{0_ free transition _ Normal pressure test}；

Step A4: in the normal-pressure wind tunnel test state, a transition tape is pasted on different chord-direction positions of the surface of the wing to carry out forced transition to obtain the gradient C of the lift line_LαAnd zero lift angle of attack alpha₀A position change rule curve along with transition; then obtaining the corresponding lifting line slope C through the step A3_{L alpha free transition normal pressure test}And zero lift angle of attack alpha_{0_ free transition _ Normal pressure test}The slope C of the lift line obtained in this step_LαAnd zero lift angle of attack alpha₀Interpolating along with the change rule curve of transition position to obtain the free transition position Xn_{Free transition normal pressure test}The position can be regarded as the free transition position corresponding to the critical Re number;

step A5: free transition position Xn obtained by step A4_{Free transition normal pressure test}Adding the difference Δ Xn obtained in step A2 to obtain a new free transition position Xn_{'forced transition' normal pressure test}The position is the corresponding free transition position under the real flight altitude and the flight M number; by bringing new free transition positionsXn_{'forced transition' normal pressure test}Slope at the lifting line C_LαAnd zero lift angle of attack alpha₀Interpolation is carried out on the curve of the position change rule along with transition to obtain a corresponding new rising line slope C_{L alpha _ forced transition _ normal pressure test}And zero lift angle of attack alpha_{0_ forced transition _ Normal pressure test}The slope C of the lifting line can be obtained_{L alpha _ forced transition _ normal pressure test}Slope C of the lift line_{L alpha free transition normal pressure test}Delta of (D) Δ C_{L alpha _ forced transition}Zero lift angle of attack alpha_{0_ forced transition _ Normal pressure test}Angle of attack alpha of zero lift_{0_ free transition _ Normal pressure test}Delta Δ α of (1)_{0_ forced transition}；

Step A6: from zero lift angle delta alpha_{0_ forced transition}Cruise angle of attack alpha and zero lift angle of attack alpha_{0_ free transition _ Normal pressure test}The difference Δ C of the slope of the lifting line_{L alpha _ forced transition}Slope of lift line C_{L alpha free transition normal pressure test}Calculating lift coefficient C_LCorrection amount of (1) (Δ C)_{L _ Compulsory transition}，

Namely, Delta C_{L _ Compulsory transition}＝-Δα_{0_ forced transition}×C_{L alpha free transition normal pressure test}+ΔC_{L alpha _ forced transition}×(α-α_{0_ free transition _ Normal pressure test})；

Step A7: superimposing the lift coefficient C of the reference state in step A1_{L _ boost}And coefficient of lift C in step A6_LCorrection amount of (1) (Δ C)_{L _ Compulsory transition}I.e. C_{L_Re}＝C_{L _ boost}+ΔC_{L _ Compulsory transition}(ii) a Obtaining the corrected lift coefficient C_{L_Re}And finishing the correction of the lift coefficient.

II, the minimum resistance coefficient C_DminThe specific steps of the correction are as follows:

step B1: increasing the proper pressure during the test to increase the Re number to the critical Re number or more and obtain the minimum resistance coefficient C of the reference state_{D _ supercharging}；

Step B2: determining corrected flight height and flight M number, calculating corresponding real Re number, and combining the change curve of laminar flow region of upper surface of wing along with Re number in real flight stateLine for obtaining the free transition position Xn corresponding to the true Re number_{Free transition (CFD)}Simultaneously obtaining the free transition position Xn corresponding to the critical Re number_{Free transition critical}To find the free transition position Xn_{Free transition (CFD)}And a free transition position Xn_{Free transition critical}The difference Δ Xn of (d);

step B3: carrying out normal-pressure wind tunnel test on the laminar flow wing type airplane to obtain the aerodynamic force characteristic of the airplane in the free transition state and obtain the corresponding minimum resistance coefficient C_{D alpha free transition normal pressure test}And zero lift angle of attack alpha_{0_ free transition _ Normal pressure test}；

Step B4: in the normal-pressure wind tunnel test state, the minimum resistance coefficient C is obtained by forcibly transition by sticking transition strips at different chord-direction positions on the surface of the wing_DαAnd zero lift angle of attack alpha₀A position change rule curve along with transition; then obtaining the corresponding minimum resistance coefficient C through the step B3_{D alpha free transition normal pressure test}And zero lift angle of attack alpha_{0_ free transition _ Normal pressure test}The minimum resistance coefficient C obtained in this step_DαAnd zero lift angle of attack alpha₀Interpolating along with the change rule curve of transition position to obtain the free transition position Xn_{Free transition normal pressure test}The position can be regarded as the free transition position corresponding to the critical Re number;

step B5: free transition position Xn obtained by step B4_{Free transition normal pressure test}Adding the difference Δ Xn obtained in step B2 to obtain a new free transition position Xn_{'forced transition' normal pressure test}The position is the corresponding free transition position under the real flight altitude and the flight M number; by bringing a new free transition position Xn_{'forced transition' normal pressure test}At the minimum drag coefficient C_DαAnd zero lift angle of attack alpha₀Interpolation is carried out on the curve of the position change rule along with transition to obtain a corresponding new minimum resistance coefficient C_{D alpha _ forced transition _ normal pressure test}And zero lift angle of attack alpha_{0_ forced transition _ Normal pressure test}The minimum resistance coefficient C can be obtained_{D alpha _ forced transition _ normal pressure test}And the minimum resistance coefficient C_{D alpha free transition normal pressure test}Delta of (D) Δ C_{D alpha _ forced transition}Zero lift angle of attack alpha_{0_ forced transition _ Normal pressure test}Angle of attack alpha of zero lift_{0_ free transition _ Normal pressure test}Delta Δ α of (1)_{0_ forced transition}；

Step B6: from zero lift angle delta alpha_{0_ forced transition}Cruise angle of attack alpha and zero lift angle of attack alpha_{0_ free transition _ Normal pressure test}The minimum resistance coefficient difference DeltaC_{D alpha _ forced transition}Minimum coefficient of resistance C_{D alpha free transition normal pressure test}Calculating the minimum resistance coefficient C_DCorrection amount of (1) (Δ C)_{D _ Compulsory transition}，

Namely, Delta C_{D _ Compulsory transition}＝-Δα_{0_ forced transition}×C_{D alpha free transition normal pressure test}+ΔC_{D alpha _ forced transition}×(α-α_{0_ free transition _ Normal pressure test})；

Step B7: superimposing the minimum drag coefficient C of the reference state in step B1_{D _ supercharging}And the minimum drag coefficient C in step B6_DCorrection amount of (1) (Δ C)_{D _ Compulsory transition}I.e. C_{D_Re}＝C_{D _ supercharging}+ΔC_{D _ Compulsory transition}(ii) a Obtaining the corrected minimum resistance coefficient C_{D_Re}And finishing the correction of the minimum resistance coefficient.

Third, the longitudinal moment coefficient C_mThe specific steps of the correction are as follows:

step C1: increasing proper pressure in the test process to increase the Re number to be tested to the critical Re number or more to the critical Re number, and acquiring the longitudinal moment coefficient C of the reference state_{m _ boost}；

Step C2: determining the corrected flight height and flight M number, calculating the corresponding real Re number, and obtaining the free transition position Xn corresponding to the real Re number by combining the change curve of the laminar flow region of the upper surface of the wing along with the Re number in the real flight state_{Free transition (CFD)}Simultaneously obtaining the free transition position Xn corresponding to the critical Re number_{Free transition critical}To find the free transition position Xn_{Free transition (CFD)}And a free transition position Xn_{Free transition critical}The difference Δ Xn of (d);

step C3: carrying out normal pressure wind tunnel test on the laminar flow wing type airplane to obtain the free transition stateThe aerodynamic characteristics of the aircraft are obtained to obtain the corresponding longitudinal moment coefficient C_{m alpha free transition normal pressure test}And zero lift moment coefficient C_{m (alpha is 0) _ free transition _ normal pressure test}；

Step C4: in the normal-pressure wind tunnel test state, the longitudinal moment coefficient C is obtained by forcibly transition by sticking transition strips at different chord-direction positions on the surface of the wing_mαAnd zero lift moment coefficient C_m(α＝0)A position change rule curve along with transition; then the corresponding longitudinal moment coefficient C obtained in the step C3 is used_{m alpha free transition normal pressure test}And zero lift moment coefficient C_{m (alpha is 0) _ free transition _ normal pressure test}The longitudinal moment coefficient C obtained in this step_mαInterpolating along with the change rule curve of transition position to obtain the free transition position Xn_{Free transition normal pressure test}The position can be regarded as the free transition position corresponding to the critical Re number;

step C5: the free transition position Xn obtained in step C4_{Free transition normal pressure test}Adding the difference Δ Xn obtained in step C2 to obtain a new free transition position Xn_{'forced transition' normal pressure test}The position is the corresponding free transition position under the real flight altitude and the flight M number; by bringing a new free transition position Xn_{'forced transition' normal pressure test}In the longitudinal direction of the moment coefficient C_mαInterpolation is carried out on the curve of the position change rule along with transition to obtain a corresponding new longitudinal moment coefficient C_{m alpha _ forced transition _ normal pressure test}And a new zero lift moment C_{m (alpha is 0) _ forced transition _ normal pressure test}The longitudinal moment coefficient C can be obtained_{m alpha _ forced transition _ normal pressure test}Coefficient of longitudinal moment C_{m alpha free transition normal pressure test}Delta of (D) Δ C_{m alpha _ forced transition}(ii) a While obtaining zero lift moment coefficient C_{m (alpha is 0) _ forced transition _ normal pressure test}Coefficient of zero lift moment C_{m (alpha is 0) _ free transition _ normal pressure test}Delta of (D) Δ C_{m (alpha 0) _ forced transition}

Step C6: from cruise angle of attack alpha, zero lift moment coefficient difference delta C_{m0_ Compulsory transition}Longitudinal moment coefficient difference Delta C_{m alpha _ forced transition}Calculating the longitudinal moment coefficient C_mRepair ofPositive quantity Δ C_{m _ Compulsory transition}，

Namely, Delta C_{m _ Compulsory transition}＝ΔC_{m (alpha 0) _ forced transition}+ΔC_{m alpha _ forced transition}×α；

Step C7: superimposing the longitudinal moment coefficient C of the reference state in step C1_{m _ boost}And the longitudinal moment coefficient C in step C6_mCorrection amount of (1) (Δ C)_{m _ Compulsory transition}I.e. C_{m_Re}＝C_{m _ boost}+ΔC_{m _ Compulsory transition}(ii) a Obtaining the corrected longitudinal moment coefficient C_{m_Re}And finishing the correction of the longitudinal moment coefficient.

A supercritical laminar flow composite wing type Re number effect correction method based on flow transition mainly comprises two steps of wing characteristic analysis and data correction; specifically, the method comprises the steps of firstly, acquiring the aerodynamic characteristics of the laminar flow wing aircraft within the full Re number range by using a CFD simulation technology, determining the critical Re of the laminar flow wing aircraft, and analyzing the wing characteristics; then, acquiring the Re number effect correction quantity of the laminar flow wing type airplane step by step through a means of combining a pressurization test and a forced transition test; and finally, finishing the Re number effect correction of the wind tunnel test data according to the Re number effect correction quantity.

Compared with the prior art, the invention has the following advantages and beneficial effects:

(1) the method not only considers the Re number to influence the flow field structure and the pressure distribution of the laminar flow airfoil, but also considers the influence of the Re number on the transition position of the laminar flow airfoil, and directly aims at the flow characteristic of the laminar flow airfoil to improve the accuracy of the Re number correction method of the laminar flow airfoil airplane.

(2) The method utilizes a CFD simulation means to obtain the aerodynamic characteristics of the laminar flow airfoil wing in the full Re number range, directly obtains the flow mechanism and the aerodynamic characteristics of the surface of the wing, and is visual and effective.

(3) According to the method, the Re number effect correction quantity of the laminar flow wing type airplane is obtained step by step through a means of combining a pressurization test and a forced transition test, and the data is real and reliable; and the main effect of high Re number is embodied through forced transition, the number of times of pressurizing wind tunnel tests is greatly reduced, and the test cost is greatly saved.

(4) The invention can be applied to various airplanes with high cruising efficiency, such as laminar flow wing profiles with high lift-drag ratio, supercritical laminar flow composite wing profiles and the like, and has wide application range.

Drawings

Fig. 1 is a schematic flow chart of a laminar flow airfoil Re number effect correction method based on flow transition.

Fig. 2 is a flight envelope of an aircraft.

FIG. 3 is a CFD simulation calculated upper surface laminar flow region Xn.jy of an airfoil_{(cruise. alpha.)}The variation curve with Re number.

FIG. 4 is a CFD simulation calculated wing lift coefficient C_{L (cruise alpha)}The variation curve with Re number.

FIG. 5 is a CFD simulation calculated wing drag coefficient C_{D (cruise alpha)}The variation curve with Re number.

FIG. 6 is a CFD simulation calculated wing lift-drag ratio K_{(cruise. alpha.)}The variation curve with Re number.

Fig. 7 is a photograph of a forced transition test.

FIG. 8 is a slope C of a lift line obtained by a forced transition test_LαA transition position relationship curve.

FIG. 9 is a minimum resistance coefficient C obtained by the forced transition test_DminA transition position relationship curve.

FIG. 10 shows a zero attack angle α obtained by a forced transition test₀A transition position relationship curve.

FIG. 11 shows a zero lift moment C obtained by the forced transition test_m(α＝0)A transition position relationship curve.

FIG. 12 is the longitudinal moment derivative C obtained by the forced transition test_mαA transition position relationship curve.

Detailed Description

The present invention will be described in further detail with reference to examples, but the embodiments of the present invention are not limited thereto.

Example 1:

the embodiment provides a laminar flow airfoil shape Re number effect correction method based on flow transition, and the method described below is executed under a certain M number, and each flight M number needs to completely execute the program.

The laminar flow airfoil Re number effect correction method based on flow transition mainly comprises two steps of wing characteristic analysis and data correction; specifically, the method comprises the steps of firstly, acquiring the aerodynamic characteristics of the laminar flow wing aircraft within the full Re number range by using a CFD simulation technology, determining the critical Re of the laminar flow wing aircraft, and analyzing the wing characteristics; then, acquiring the Re number effect correction quantity of the laminar flow wing type airplane step by step through a means of combining a pressurization test and a forced transition test; and finally, finishing the Re number effect correction of the wind tunnel test data according to the Re number effect correction quantity.

A flow chart of a laminar flow airfoil Re number effect correction method based on flow transition is shown in fig. 1, and includes the following steps:

step S100: analyzing the characteristics of the wing;

step S200: and (6) data correction.

The wing characteristic analysis of step S100 specifically includes the steps of:

step S110: determining the flight Re number range of the airplane according to the flight envelope of the airplane, as shown in figure 2;

step S120: calculating aerodynamic characteristics of laminar flow wings within the flight Re number range by using a CFD simulation technology to obtain the flow area Xn.jy of the upper surface layer of the wings in a real flight state_{(cruise. alpha.)}Variation curve along with Re number and wing lift coefficient C_{L (cruise alpha)}Variation curve along with Re number and wing resistance coefficient C_{D (cruise alpha)}Variation curve along with Re number and wing lift-drag ratio K_{(cruise. alpha.)}The variation curve with the number of Re;

wherein, the surface flow area Xn.jy on the wing under the real flight state_{(cruise. alpha.)}The variation curve with Re is shown in FIG. 3;

coefficient of lift of wing C under real flight state_{L (cruise alpha)}The variation curve with Re is shown in FIG. 4;

coefficient of wing drag C under true flight conditions_{D (cruise alpha)}The variation curve with Re is shown in FIG. 5;

wing lift-drag ratio K under real flight state_{(cruise. alpha.)}The variation curve with Re is shown in FIG. 6;

step S130: each change curve in step S120 is observed, and a critical Re number is selected from the plurality of Re numbers.

It can be seen from the observation of the curves that below a certain number Re, the laminar flow area of the upper surface of the wing remains substantially unchanged, but the aerodynamic performance of the wing is poor. Along with the increase of the Re number, diffusion separation and diffusion-shock wave separation in the airflow boundary layer are gradually eliminated, the lift force is increased, the resistance is reduced, and the lift-drag ratio is improved. When the number of the Re is larger than a certain number, the laminar flow area is reduced along with the increase of the number of Re, the laminar flow transition point gradually moves forwards, the lift force is reduced, the resistance is increased, the lift-drag ratio is reduced, and finally the upper surface of the wing is changed into a full turbulent flow state.

That is, as shown in fig. 3, the critical Re number satisfies the following three conditions at the same time:

condition a: when the number of the laminar flow areas is less than the critical Re number, the laminar flow area of the upper surface of the wing basically keeps unchanged, but the aerodynamic performance of the wing is poor;

condition b: along with the increase of the Re number, diffusion separation and diffusion-shock wave separation in the airflow boundary layer are gradually eliminated, the lift force is increased, the resistance is reduced, and the lift-drag ratio is improved;

condition c: after the number of the Re is larger than the critical Re number, the laminar flow area is reduced along with the increase of the Re number, the laminar flow transition point gradually moves forwards, the lift force is reduced, the resistance is increased, the lift-drag ratio is reduced, and finally the upper surface of the wing is changed into a full turbulent flow state.

The step S200 performs data correction. In the data correction process, firstly, the real flying Re number corresponding to the real flying height and the flying M number is judged according to the flying envelope of the airplane, and then one of the following modes is selected according to the magnitude relation between the real flying Re number and the critical Re number to obtain the Re number effect correction quantity of the laminar flow wing type airplane

The first method is as follows: if the real flight Re number is smaller than the critical Re number, the test Re number is increased to the real flight Re number or is closer to the real flight Re number by adopting a pressurization test during the test, and the laminar flow wing type airplane Re number effect correction amount in the test data is directly obtained;

the second method comprises the following steps: if the real flying Re number is larger than or equal to the critical Re number, increasing the test Re number to the critical Re number or closer to the critical Re by adopting a pressurization test, and taking the data as reference data; carrying out a forced transition test by manually pasting transition belts at different chord-wise positions on the surface of the wing, obtaining lift force, resistance and longitudinal torque increment caused by transition position change, and superposing the lift force, the resistance and the longitudinal torque increment on reference data to obtain Re number effect correction of the laminar flow wing type airplane in the test data;

in which, a transition tape is pasted on different chord-wise positions on the surface of the wing for a forced transition test, as shown in fig. 7.

That is, when the data correction is performed in step S200, there are two ways, one way and two ways, depending on the relationship between the critical Re number and the Re number corresponding to the real flying height and the flying M number.

The embodiment combines the contents of the embodiment 2, the embodiment 3 and the embodiment 4, is successfully applied to the Re number effect correction of the two-type unmanned aerial vehicle, and proves the rationality and the accuracy of the laminar flow airfoil Re number effect correction method based on the flow transition through the trial flight verification.

Example 2:

the embodiment is further optimized on the basis of the embodiment 1, and the Re number effect correction comprises a lift coefficient C_LModified, minimum drag coefficient C_DminCorrection, longitudinal moment coefficient C_mAnd (6) correcting.

First, with the lift coefficient C of the wing_LThe correction is taken as an example, and the specific steps are as follows:

step A1: increasing proper pressure in the test process to increase the Re number of the test to the critical Re number or more to the critical Re number, and acquiring the lift coefficient C of the reference state_{L _ boost}；

Step A2: determining the corrected flying height and flying M number, calculating the corresponding real Re number, and obtaining the free transition position Xn corresponding to the real Re number by combining the change curve of the upper surface laminar flow region of the wing along with the Re number in the real flying state as shown in FIG. 3_{Free transition (CFD)}Simultaneously obtaining the free transition position Xn corresponding to the critical Re number_{Free transition critical}To find the free transition position Xn_{Free transition (CFD)}Position X of transition with freedomn_{Free transition critical}The difference Δ Xn of (d);

step A3: carrying out normal-pressure wind tunnel test on the laminar flow wing type airplane to obtain the aerodynamic force characteristic of the airplane in the free transition state and obtain the corresponding lift line slope C_{L alpha free transition normal pressure test}And zero lift angle of attack alpha_{0_ free transition _ Normal pressure test}；

Step A4: in the normal pressure wind tunnel test state, a transition tape is pasted on different chord direction positions on the surface of the wing to carry out forced transition to obtain the gradient C of the lift line shown in fig. 8 and 10_LαAnd zero lift angle of attack alpha₀A position change rule curve along with transition; then obtaining the corresponding lifting line slope C through the step A3_{L alpha free transition normal pressure test}And zero lift angle of attack alpha_{0_ free transition _ Normal pressure test}The slope C of the lift line obtained in this step_LαAnd zero lift angle of attack alpha₀Interpolating along with the change rule curve of transition position to obtain the free transition position Xn_{Free transition normal pressure test}The position can be regarded as the free transition position corresponding to the critical Re number;

step A5: free transition position Xn obtained by step A4_{Free transition normal pressure test}Adding the difference Δ Xn obtained in step A2 to obtain a new free transition position Xn_{'forced transition' normal pressure test}The position is the corresponding free transition position under the real flight altitude and the flight M number; by bringing a new free transition position Xn_{'forced transition' normal pressure test}Slope at the lifting line C_LαAnd zero lift angle of attack alpha₀Interpolation is performed on the curve of the transition position change rule to obtain a new slope C of the lift line as shown in fig. 8 and 10_{L alpha _ forced transition _ normal pressure test}And zero lift angle of attack alpha_{0_ forced transition _ Normal pressure test}The slope C of the lifting line can be obtained_{L alpha _ forced transition _ normal pressure test}Slope C of the lift line_{L alpha free transition normal pressure test}Delta of (D) Δ C_{L alpha _ forced transition}Zero lift angle of attack alpha_{0_ forced transition _ Normal pressure test}Angle of attack alpha of zero lift_{0_ free transition _ Normal pressure test}Delta Δ α of (1)_{0_ forced transition}；

Step A6: angle of attack from zero liftDelta alpha_{0_ forced transition}Cruise angle of attack alpha and zero lift angle of attack alpha_{0_ free transition _ Normal pressure test}The difference Δ C of the slope of the lifting line_{L alpha _ forced transition}Slope of lift line C_{L alpha free transition normal pressure test}Calculating a lift coefficient C by the formula (1)_LCorrection amount of (1) (Δ C)_{L _ Compulsory transition}；

ΔC_{L _ Compulsory transition}＝-Δα_{0_ forced transition}×C_{L alpha free transition normal pressure test}+ΔC_{L alpha _ forced transition}×(α-α_{0_ free transition _ Normal pressure test}) Formula (1);

step A7: superimposing the lift coefficient C of the reference state in step A1 by equation (2)_{L _ boost}And coefficient of lift C in step A6_LCorrection amount of (1) (Δ C)_{L _ Compulsory transition}Obtaining a corrected lift coefficient C_{L_Re}Finishing the correction of the lift coefficient;

C_{L_Re}＝C_{l _ boost}+ΔC_{L _ Compulsory transition}Equation (2).

Other parts of this embodiment are the same as embodiment 1, and thus are not described again.

Example 3:

similarly, the step in embodiment 2 can be used for correcting the drag coefficient of the wing, and the drag coefficient of the wing is obtained by correcting the drag coefficient of the wing by using the formula (3);

C_{Dmin_Re}＝C_{dmin _ boost}+ΔC_{Dmin _ forced transition}Equation (3).

Specifically, step B1: increasing the proper pressure during the test to increase the Re number to the critical Re number or more and obtain the minimum resistance coefficient C of the reference state_{D _ supercharging}；

Step B2: determining the corrected flight height and flight M number, calculating the corresponding real Re number, and obtaining the free transition position Xn corresponding to the real Re number by combining the change curve of the laminar flow region of the upper surface of the wing along with the Re number in the real flight state_{Free transition (CFD)}Simultaneously obtaining the free transition position Xn corresponding to the critical Re number_{Free transition critical}To find the free transition position Xn_{Free transition (CFD)}And a free transition position Xn_{Free transition critical}The difference Δ Xn of (d);

step B3: carrying out normal-pressure wind tunnel test on the laminar flow wing type airplane to obtain the aerodynamic force characteristic of the airplane in the free transition state and obtain the corresponding minimum resistance coefficient C_{D alpha free transition normal pressure test}And zero lift angle of attack alpha_{0_ free transition _ Normal pressure test}；

Step B4: in the normal-pressure wind tunnel test state, the minimum resistance coefficient C is obtained by forcibly transition by sticking transition strips at different chord-direction positions on the surface of the wing_DαAnd zero lift angle of attack alpha₀A position change rule curve along with transition; then obtaining the corresponding minimum resistance coefficient C through the step B3_{D alpha free transition normal pressure test}And zero lift angle of attack alpha_{0_ free transition _ Normal pressure test}The minimum resistance coefficient C obtained in this step_DαAnd zero lift angle of attack alpha₀Interpolating along with the change rule curve of transition position to obtain the free transition position Xn_{Free transition normal pressure test}The position can be regarded as the free transition position corresponding to the critical Re number;

step B5: free transition position Xn obtained by step B4_{Free transition normal pressure test}Adding the difference Δ Xn obtained in step B2 to obtain a new free transition position Xn_{'forced transition' normal pressure test}The position is the corresponding free transition position under the real flight altitude and the flight M number; by bringing a new free transition position Xn_{'forced transition' normal pressure test}At the minimum drag coefficient C_DαAnd zero lift angle of attack alpha₀Interpolation is carried out on the curve of the position change rule along with the transition to obtain a corresponding new minimum resistance coefficient C as shown in fig. 9 and 10_{D alpha _ forced transition _ normal pressure test}And zero lift angle of attack alpha_{0_ forced transition _ Normal pressure test}The minimum resistance coefficient C can be obtained_{D alpha _ forced transition _ normal pressure test}And the minimum resistance coefficient C_{D alpha free transition normal pressure test}Delta of (D) Δ C_{D alpha _ forced transition}Zero lift angle of attack alpha_{0_ forced transition _ Normal pressure test}Angle of attack alpha of zero lift_{0_ free transition _ Normal pressure test}Delta Δ α of (1)_{0_ forced transition}；

Step B6: from zero lift angle differenceQuantity Δ α_{0_ forced transition}Cruise angle of attack alpha and zero lift angle of attack alpha_{0_ free transition _ Normal pressure test}The minimum resistance coefficient difference DeltaC_{D alpha _ forced transition}Minimum coefficient of resistance C_{D alpha free transition normal pressure test}Calculating the minimum resistance coefficient C_DCorrection amount of (1) (Δ C)_{D _ Compulsory transition}，

Namely, Delta C_{D _ Compulsory transition}＝-Δα_{0_ forced transition}×C_{D alpha free transition normal pressure test}+ΔC_{D alpha _ forced transition}×(α-α_{0_ free transition _ Normal pressure test})；

Step B7: superimposing the minimum drag coefficient C of the reference state in step B1_{D _ supercharging}And the minimum drag coefficient C in step B6_DCorrection amount of (1) (Δ C)_{D _ Compulsory transition}I.e. C_{D_Re}＝C_{D _ supercharging}+ΔC_{D _ Compulsory transition}(ii) a Obtaining the corrected minimum resistance coefficient C_{D_Re}And finishing the correction of the minimum resistance coefficient.

Other parts of this embodiment are the same as embodiment 1, and thus are not described again.

Example 4:

similarly, longitudinal moment coefficient C_mThe step in the embodiment 2 or 3 can be used for obtaining the longitudinal moment coefficient C by using the formula (4) and the formula (5) for correction_{m_Re}。

C_{m_Re}＝C_{m _ boost}+ΔC_{m _ Compulsory transition}Formula (4);

ΔC_{m _ Compulsory transition}＝ΔC_{m (alpha 0) _ forced transition}+ΔC_{m alpha _ forced transition}XAlpha equation (5).

In particular, the longitudinal moment coefficient C_mThe specific steps of the correction are as follows:

step C1: increasing proper pressure in the test process to increase the Re number to be tested to the critical Re number or more to the critical Re number, and acquiring the longitudinal moment coefficient C of the reference state_{m _ boost}；

Step C2: determining corrected flight height and flight M number, calculating corresponding real Re number, and combining the change curve of the laminar flow region of the upper surface of the wing along with the Re number in the real flight stateObtaining the free transition position Xn corresponding to the true Re number_{Free transition (CFD)}Simultaneously obtaining the free transition position Xn corresponding to the critical Re number_{Free transition critical}To find the free transition position Xn_{Free transition (CFD)}And a free transition position Xn_{Free transition critical}The difference Δ Xn of (d);

step C3: carrying out normal-pressure wind tunnel test on the laminar flow wing type airplane to obtain the aerodynamic force characteristic of the airplane in the free transition state, and obtaining the corresponding longitudinal moment coefficient C_{m alpha free transition normal pressure test}And zero lift moment coefficient C_{m (alpha is 0) _ free transition _ normal pressure test}；

Step C4: in the normal-pressure wind tunnel test state, the longitudinal moment coefficient C is obtained by forcibly transition by sticking transition strips at different chord-direction positions on the surface of the wing_mαAnd zero lift moment coefficient C_m(α＝0)A position change rule curve along with transition; then the corresponding longitudinal moment coefficient C obtained in the step C3 is used_{m alpha free transition normal pressure test}And zero lift moment coefficient C_{m (alpha is 0) _ free transition _ normal pressure test}The longitudinal moment coefficient C obtained in this step_mαInterpolating along with the change rule curve of transition position to obtain the free transition position Xn_{Free transition normal pressure test}The position can be regarded as the free transition position corresponding to the critical Re number;

step C5: the free transition position Xn obtained in step C4_{Free transition normal pressure test}Adding the difference Δ Xn obtained in step C2 to obtain a new free transition position Xn_{'forced transition' normal pressure test}The position is the corresponding free transition position under the real flight altitude and the flight M number; by bringing a new free transition position Xn_{'forced transition' normal pressure test}In the longitudinal direction of the moment coefficient C_mαInterpolation is performed on the curve of the position change rule along with the transition to obtain a corresponding new longitudinal moment coefficient C as shown in fig. 11 and 12_{m alpha _ forced transition _ normal pressure test}And a new zero lift moment C_{m (alpha is 0) _ forced transition _ normal pressure test}The longitudinal moment coefficient C can be obtained_{m alpha _ forced transition _ normal pressure test}Coefficient of longitudinal moment C_{m alpha free transition normal pressure test}Delta of (D) Δ C_{m alpha _ forced transition}(ii) a While obtaining zero lift moment coefficient C_{m (alpha is 0) _ forced transition _ normal pressure test}Coefficient of zero lift moment C_{m (alpha is 0) _ free transition _ normal pressure test}Delta of (D) Δ C_{m (alpha 0) _ forced transition}

Step C6: from cruise angle of attack alpha, zero lift moment coefficient difference delta C_{m (alpha 0) _ forced transition}Longitudinal moment coefficient difference Delta C_{m alpha _ forced transition}Calculating the longitudinal moment coefficient C_mCorrection amount of (1) (Δ C)_{m _ Compulsory transition}，

Namely, Delta C_{m _ Compulsory transition}＝ΔC_{m (alpha 0) _ forced transition}+ΔC_{m alpha _ forced transition}×α；

Step C7: superimposing the longitudinal moment coefficient C of the reference state in step C1_{m _ boost}And the longitudinal moment coefficient C in step C6_mCorrection amount of (1) (Δ C)_{m _ Compulsory transition}I.e. C_{m_Re}＝C_{m _ boost}+ΔC_{m _ Compulsory transition}(ii) a Obtaining the corrected longitudinal moment coefficient C_{m_Re}And finishing the correction of the longitudinal moment coefficient.

Other parts of this embodiment are the same as those of embodiment 1, embodiment 2, or embodiment 3, and thus are not described again.

Example 5:

the embodiment is optimized on the basis of any one of embodiments 1 to 4, and the Re number effect correction method is applied to the supercritical laminar flow composite wing aircraft.

A supercritical laminar flow composite wing type Re number effect correction method based on flow transition mainly comprises two steps of wing characteristic analysis and data correction; specifically, the method comprises the steps of firstly, acquiring the aerodynamic characteristics of the laminar flow wing aircraft within the full Re number range by using a CFD simulation technology, determining the critical Re of the laminar flow wing aircraft, and analyzing the wing characteristics; then, acquiring the Re number effect correction quantity of the laminar flow wing type airplane step by step through a means of combining a pressurization test and a forced transition test; and finally, finishing the Re number effect correction of the wind tunnel test data according to the Re number effect correction quantity.

Other parts of this embodiment are the same as any of embodiments 1 to 4, and thus are not described again.

The above description is only a preferred embodiment of the present invention, and is not intended to limit the present invention in any way, and all simple modifications and equivalent variations of the above embodiments according to the technical spirit of the present invention are included in the scope of the present invention.

Claims (10)

1. A laminar flow airfoil type Re number effect correction method based on flow transition is characterized by mainly comprising two steps of wing characteristic analysis and data correction; specifically, the method comprises the steps of firstly, acquiring the aerodynamic characteristics of the laminar flow wing aircraft within the full Re number range by using a CFD simulation technology, determining the critical Re of the laminar flow wing aircraft, and analyzing the wing characteristics; then, acquiring the Re number effect correction quantity of the laminar flow wing type airplane step by step through a means of combining a pressurization test and a forced transition test; finally, finishing the Re number effect correction of the wind tunnel test data according to the Re number effect correction quantity;

the wing characteristic analysis specifically comprises the following steps:

step S110: calculating aerodynamic characteristics of laminar flow wings within the flight Re number range by using a CFD simulation technology to obtain the flow area Xn.jy of the upper surface layer of the wings in a real flight state_{(cruise. alpha.)}Variation curve along with Re number, wing lift coefficient C_{L (cruise alpha)}Variation curve along with Re number and wing resistance coefficient C_{D (cruise alpha)}Variation curve along with Re number and wing lift-drag ratio K_{(cruise. alpha.)}The variation curve with the number of Re;

step S120: observing each change curve in the step S110, and selecting a critical Re number from a plurality of Re numbers;

wherein the critical Re number simultaneously satisfies the following three conditions:

condition a: when the number of the laminar flow areas is less than the critical Re number, the laminar flow area of the upper surface of the wing basically keeps unchanged, but the aerodynamic performance of the wing is poor;

condition b: along with the increase of the Re number, diffusion separation and diffusion-shock wave separation in the airflow boundary layer are gradually eliminated, the lift force is increased, the resistance is reduced, and the lift-drag ratio is improved;

condition c: after the number of the Re is larger than the critical Re number, the laminar flow area is reduced along with the increase of the Re number, the laminar flow transition point gradually moves forwards, the lift force is reduced, the resistance is increased, the lift-drag ratio is reduced, and finally the upper surface of the wing is changed into a full turbulent flow state;

in the data correction process, firstly, the real flying Re number corresponding to the real flying height and the flying M number is judged according to the flying envelope of the airplane, and then one of the following modes is selected according to the magnitude relation between the real flying Re number and the critical Re number to obtain the Re number effect correction quantity of the laminar flow airfoil airplane:

the first method is as follows: if the real flight Re number is smaller than the critical Re number, the test Re number is increased to the real flight Re number or is closer to the real flight Re number by adopting a pressurization test during the test, and the laminar flow wing type airplane Re number effect correction amount in the test data is directly obtained;

the second method comprises the following steps: if the real flying Re number is larger than or equal to the critical Re number, increasing the test Re number to the critical Re number or closer to the critical Re by adopting a pressurization test, and taking the data as reference data; and performing a forced transition test by manually pasting transition belts at different chord-wise positions on the surface of the wing to obtain lift force, resistance and longitudinal torque increment caused by transition position change, and superposing the lift force, the resistance and the longitudinal torque increment on reference data to obtain Re number effect correction of the laminar flow wing type airplane in the test data.

2. The method for modifying the Re effect of the laminar airfoil based on flow transition as claimed in claim 1, wherein: the Re number effect correction comprises a lift coefficient C_LModified, minimum drag coefficient C_DminCorrection, longitudinal moment coefficient C_mAnd (6) correcting.

3. The method for modifying the Re effect of the laminar airfoil based on flow transition as claimed in claim 2, wherein: coefficient of lift C_LThe specific steps of the correction are as follows:

step A1: increasing proper pressure in the test process to increase the Re number of the test to the critical Re number or more to the critical Re number, and acquiring the lift coefficient C of the reference state_{L _ boost}；

Step A2: determining corrected flying height and flying M number, calculating corresponding real Re number, and combining with trueObtaining a free transition position Xn corresponding to the real Re number according to the change curve of the upper surface laminar flow region of the wing along with the Re number in the real flight state_{Free transition (CFD)}Simultaneously obtaining the free transition position Xn corresponding to the critical Re number_{Free transition critical}To find the free transition position Xn_{Free transition (CFD)}And a free transition position Xn_{Free transition critical}The difference Δ Xn of (d);

step A3: carrying out normal-pressure wind tunnel test on the laminar flow wing type airplane to obtain the aerodynamic force characteristic of the airplane in the free transition state and obtain the corresponding lift line slope C_{L alpha free transition normal pressure test}And zero lift angle of attack alpha_{0_ free transition _ Normal pressure test}；

Step A4: in the normal-pressure wind tunnel test state, a transition tape is pasted on different chord-direction positions of the surface of the wing to carry out forced transition to obtain the gradient C of the lift line_LαAnd zero lift angle of attack alpha₀A position change rule curve along with transition; then obtaining the corresponding lifting line slope C through the step A3_{L alpha free transition normal pressure test}And zero lift angle of attack alpha_{0_ free transition _ Normal pressure test}The slope C of the lift line obtained in this step_LαAnd zero lift angle of attack alpha₀Interpolating along with the change rule curve of transition position to obtain the free transition position Xn_{Free transition normal pressure test}The position can be regarded as the free transition position corresponding to the critical Re number;

step A5: free transition position Xn obtained by step A4_{Free transition normal pressure test}Adding the difference Δ Xn obtained in step A2 to obtain a new free transition position Xn_{'forced transition' normal pressure test}The position is the corresponding free transition position under the real flight altitude and the flight M number; by bringing a new free transition position Xn_{'forced transition' normal pressure test}Slope at the lifting line C_LαAnd zero lift angle of attack alpha₀Interpolation is carried out on the curve of the position change rule along with transition to obtain a corresponding new rising line slope C_{L alpha _ forced transition _ normal pressure test}And zero lift angle of attack alpha_{0_ forced transition _ Normal pressure test}The slope C of the lifting line can be obtained_{L alpha _ forced transition _ normal pressure test}Slope C of the lift line_{L alpha free transition normal pressure test}Delta of (D) Δ C_{L alpha _ forced transition}Zero lift angle of attack alpha_{0_ forced transition _ Normal pressure test}Angle of attack alpha of zero lift_{0_ free transition _ Normal pressure test}Delta Δ α of (1)_{0_ forced transition}；

Step A6: from zero lift angle delta alpha_{0_ forced transition}Cruise angle of attack alpha and zero lift angle of attack alpha_{0_ free transition _ Normal pressure test}The difference Δ C of the slope of the lifting line_{L alpha _ forced transition}Slope of lift line C_{L alpha free transition normal pressure test}Calculating lift coefficient C_LCorrection amount of (1) (Δ C)_{L _ Compulsory transition}I.e. Δ C_{L _ Compulsory transition}＝-Δα_{0_ forced transition}×C_{L alpha free transition normal pressure test}+ΔC_{L alpha _ forced transition}×(α-α_{0_ free transition _ Normal pressure test})；

Step A7: superimposing the lift coefficient C of the reference state in step A1_{L _ boost}And coefficient of lift C in step A6_LCorrection amount of (1) (Δ C)_{L _ Compulsory transition}I.e. C_{L_Re}＝C_{L _ boost}+ΔC_{L _ Compulsory transition}(ii) a Obtaining the corrected lift coefficient C_{L_Re}And finishing the correction of the lift coefficient.

4. The method for modifying the Re effect of the laminar airfoil based on flow transition as claimed in claim 2, wherein: the minimum coefficient of resistance C_DminThe specific steps of the correction are as follows:

step B1: increasing the proper pressure during the test to increase the Re number to the critical Re number or more and obtain the minimum resistance coefficient C of the reference state_{D _ supercharging}；

Step B2: determining the corrected flight height and flight M number, calculating the corresponding real Re number, and obtaining the free transition position Xn corresponding to the real Re number by combining the change curve of the laminar flow region of the upper surface of the wing along with the Re number in the real flight state_{Free transition (CFD)}Simultaneously obtaining the free transition position Xn corresponding to the critical Re number_{Free transition critical}To find the free transition position Xn_{Free transition (CFD)}And a free transition position Xn_{Free transition critical}The difference Δ Xn of (d);

step B3: carrying out normal-pressure wind tunnel test on the laminar flow wing type airplane to obtain the aerodynamic force characteristic of the airplane in the free transition state and obtain the corresponding minimum resistance coefficient C_{D alpha free transition normal pressure test}And zero lift angle of attack alpha_{0_ free transition _ Normal pressure test}；

Step B4: in the normal-pressure wind tunnel test state, the minimum resistance coefficient C is obtained by forcibly transition by sticking transition strips at different chord-direction positions on the surface of the wing_DαAnd zero lift angle of attack alpha₀A position change rule curve along with transition; then obtaining the corresponding minimum resistance coefficient C through the step B3_{D alpha free transition normal pressure test}And zero lift angle of attack alpha_{0_ free transition _ Normal pressure test}The minimum resistance coefficient C obtained in this step_DαAnd zero lift angle of attack alpha₀Interpolating along with the change rule curve of transition position to obtain the free transition position Xn_{Free transition normal pressure test}The position can be regarded as the free transition position corresponding to the critical Re number;

step B5: free transition position Xn obtained by step B4_{Free transition normal pressure test}Adding the difference Δ Xn obtained in step B2 to obtain a new free transition position Xn_{'forced transition' normal pressure test}The position is the corresponding free transition position under the real flight altitude and the flight M number; by bringing a new free transition position Xn_{'forced transition' normal pressure test}At the minimum drag coefficient C_DαAnd zero lift angle of attack alpha₀Interpolation is carried out on the curve of the position change rule along with transition to obtain a corresponding new minimum resistance coefficient C_{D alpha _ forced transition _ normal pressure test}And zero lift angle of attack alpha_{0_ forced transition _ Normal pressure test}The minimum resistance coefficient C can be obtained_{D alpha _ forced transition _ normal pressure test}And the minimum resistance coefficient C_{D alpha free transition normal pressure test}Delta of (D) Δ C_{D alpha _ forced transition}Zero lift angle of attack alpha_{0_ forced transition _ Normal pressure test}Angle of attack alpha of zero lift_{0_ free transition _ Normal pressure test}Delta Δ α of (1)_{0_ forced transition}；

Step B6: from zero lift angle delta alpha_{0_ forced transition}Cruise angle of attack alpha and zero lift angle of attack alpha_{0_ freeTransition _ Normal pressure test}The minimum resistance coefficient difference DeltaC_{D alpha _ forced transition}Minimum coefficient of resistance C_{D alpha free transition normal pressure test}Calculating the minimum resistance coefficient C_DCorrection amount of (1) (Δ C)_{D _ Compulsory transition}I.e. Δ C_{D _ Compulsory transition}＝-Δα_{0_ forced transition}×C_{D alpha free transition normal pressure test}+ΔC_{D alpha _ forced transition}×(α-α_{0_ free transition _ Normal pressure test})；

Step B7: superimposing the minimum drag coefficient C of the reference state in step B1_{D _ supercharging}And the minimum drag coefficient C in step B6_DCorrection amount of (1) (Δ C)_{D _ Compulsory transition}I.e. C_{D_Re}＝C_{D _ supercharging}+ΔC_{D _ Compulsory transition}(ii) a Obtaining the corrected minimum resistance coefficient C_{D_Re}And finishing the correction of the minimum resistance coefficient.

5. The method for modifying the Re effect of the laminar airfoil based on flow transition as claimed in claim 2, wherein: the longitudinal moment coefficient C_mThe specific steps of the correction are as follows:

step C1: increasing proper pressure in the test process to increase the Re number to be tested to the critical Re number or more to the critical Re number, and acquiring the longitudinal moment coefficient C of the reference state_{m _ boost}；

Step C2: determining the corrected flight height and flight M number, calculating the corresponding real Re number, and obtaining the free transition position Xn corresponding to the real Re number by combining the change curve of the laminar flow region of the upper surface of the wing along with the Re number in the real flight state_{Free transition (CFD)}Simultaneously obtaining the free transition position Xn corresponding to the critical Re number_{Free transition critical}To find the free transition position Xn_{Free transition (CFD)}And a free transition position Xn_{Free transition critical}The difference Δ Xn of (d);

step C3: carrying out normal-pressure wind tunnel test on the laminar flow wing type airplane to obtain the aerodynamic force characteristic of the airplane in the free transition state, and obtaining the corresponding longitudinal moment coefficient C_{m alpha free transition normal pressure test}And zero lift moment coefficient C_{m (alpha is 0) _ free transition _ normal pressure test}；

Step C4: in the normal-pressure wind tunnel test state, the longitudinal moment coefficient C is obtained by forcibly transition by sticking transition strips at different chord-direction positions on the surface of the wing_mαAnd zero lift moment coefficient C_m(α＝0)A position change rule curve along with transition; then the corresponding longitudinal moment coefficient C obtained in the step C3 is used_{m alpha free transition normal pressure test}And zero lift moment coefficient C_{m (alpha is 0) _ free transition _ normal pressure test}The longitudinal moment coefficient C obtained in this step_mαInterpolating along with the change rule curve of transition position to obtain the free transition position Xn_{Free transition normal pressure test}The position can be regarded as the free transition position corresponding to the critical Re number;

step C5: the free transition position Xn obtained in step C4_{Free transition normal pressure test}Adding the difference Δ Xn obtained in step C2 to obtain a new free transition position Xn_{'forced transition' normal pressure test}The position is the corresponding free transition position under the real flight altitude and the flight M number; by bringing a new free transition position Xn_{'forced transition' normal pressure test}In the longitudinal direction of the moment coefficient C_mαInterpolation is carried out on the curve of the position change rule along with transition to obtain a corresponding new longitudinal moment coefficient C_{m alpha _ forced transition _ normal pressure test}And a new zero lift moment C_{m (alpha is 0) _ forced transition _ normal pressure test}The longitudinal moment coefficient C can be obtained_{m alpha _ forced transition _ normal pressure test}Coefficient of longitudinal moment C_{m alpha free transition normal pressure test}Delta of (D) Δ C_{m alpha _ forced transition}(ii) a While obtaining zero lift moment coefficient C_{m (alpha is 0) _ forced transition _ normal pressure test}Coefficient of zero lift moment C_{m (alpha is 0) _ free transition _ normal pressure test}Delta of (D) Δ C_{m (alpha 0) _ forced transition}；

Step C6: from cruise angle of attack alpha, zero lift moment coefficient difference delta C_{m (alpha 0) _ forced transition}Longitudinal moment coefficient difference Delta C_{m alpha _ forced transition}Calculating the longitudinal moment coefficient C_mCorrection amount of (1) (Δ C)_{m _ Compulsory transition}I.e. Δ C_{m _ Compulsory transition}＝ΔC_{m (alpha 0) _ forced transition}+ΔC_{m alpha _ forced transition}×α；

Step C7: superimposing the longitudinal moment coefficient C of the reference state in step C1_{m _ boost}And the longitudinal moment coefficient C in step C6_mCorrection amount of (1) (Δ C)_{m _ Compulsory transition}I.e. C_{m_Re}＝C_{m _ boost}+ΔC_{m _ Compulsory transition}(ii) a Obtaining the corrected longitudinal moment coefficient C_{m_Re}And finishing the correction of the longitudinal moment coefficient.

6. A supercritical laminar flow composite wing type Re number effect correction method based on flow transition is characterized by mainly comprising two steps of wing characteristic analysis and data correction; specifically, the method comprises the steps of firstly, acquiring the aerodynamic characteristics of the laminar flow wing aircraft within the full Re number range by using a CFD simulation technology, determining the critical Re of the laminar flow wing aircraft, and analyzing the wing characteristics; then, acquiring the Re number effect correction quantity of the laminar flow wing type airplane step by step through a means of combining a pressurization test and a forced transition test; finally, finishing the Re number effect correction of the wind tunnel test data according to the Re number effect correction quantity;

the wing characteristic analysis specifically comprises the following steps:

step S110: calculating aerodynamic characteristics of laminar flow wings within the flight Re number range by using a CFD simulation technology to obtain the flow area Xn.jy of the upper surface layer of the wings in a real flight state_{(cruise. alpha.)}Variation curve along with Re number, wing lift coefficient C_{L (cruise alpha)}Variation curve along with Re number and wing resistance coefficient C_{D (cruise alpha)}Variation curve along with Re number and wing lift-drag ratio K_{(cruise. alpha.)}The variation curve with the number of Re;

step S120: observing each change curve in the step S110, and selecting a critical Re number from a plurality of Re numbers;

wherein the critical Re number simultaneously satisfies the following three conditions:

condition a: when the number of the laminar flow areas is less than the critical Re number, the laminar flow area of the upper surface of the wing basically keeps unchanged, but the aerodynamic performance of the wing is poor;

condition b: along with the increase of the Re number, diffusion separation and diffusion-shock wave separation in the airflow boundary layer are gradually eliminated, the lift force is increased, the resistance is reduced, and the lift-drag ratio is improved;

condition c: after the number of the Re is larger than the critical Re number, the laminar flow area is reduced along with the increase of the Re number, the laminar flow transition point gradually moves forwards, the lift force is reduced, the resistance is increased, the lift-drag ratio is reduced, and finally the upper surface of the wing is changed into a full turbulent flow state;

in the data correction process, firstly, the real flying Re number corresponding to the real flying height and the flying M number is judged according to the flying envelope of the airplane, and then one of the following modes is selected according to the magnitude relation between the real flying Re number and the critical Re number to obtain the Re number effect correction quantity of the laminar flow airfoil airplane:

the first method is as follows: if the real flight Re number is smaller than the critical Re number, the test Re number is increased to the real flight Re number or is closer to the real flight Re number by adopting a pressurization test during the test, and the laminar flow wing type airplane Re number effect correction amount in the test data is directly obtained;

the second method comprises the following steps: if the real flying Re number is larger than or equal to the critical Re number, increasing the test Re number to the critical Re number or closer to the critical Re by adopting a pressurization test, and taking the data as reference data; and performing a forced transition test by manually pasting transition belts at different chord-wise positions on the surface of the wing to obtain lift force, resistance and longitudinal torque increment caused by transition position change, and superposing the lift force, the resistance and the longitudinal torque increment on reference data to obtain Re number effect correction of the laminar flow wing type airplane in the test data.

7. The method for modifying the Re number effect of the flow transition-based supercritical laminar flow composite airfoil, as claimed in claim 6, wherein: the Re number effect correction comprises a lift coefficient C_LModified, minimum drag coefficient C_DminCorrection, longitudinal moment coefficient C_mAnd (6) correcting.

8. The method for modifying the Re effect of the flow transition-based supercritical laminar flow composite airfoil profile according to claim 7, wherein: coefficient of lift C_LThe specific steps of the correction are as follows:

step A1: increasing proper pressure in the test process to increase the Re number of the test to the critical Re number or more to the critical Re number, and acquiring the lift coefficient C of the reference state_{L _ boost}；

Step A2: determining the corrected flight height and flight M number, calculating the corresponding real Re number, and obtaining the free transition position Xn corresponding to the real Re number by combining the change curve of the laminar flow region of the upper surface of the wing along with the Re number in the real flight state_{Free transition (CFD)}Simultaneously obtaining the free transition position Xn corresponding to the critical Re number_{Free transition critical}To find the free transition position Xn_{Free transition (CFD)}And a free transition position Xn_{Free transition critical}The difference Δ Xn of (d);

step A3: carrying out normal-pressure wind tunnel test on the laminar flow wing type airplane to obtain the aerodynamic force characteristic of the airplane in the free transition state and obtain the corresponding lift line slope C_{L alpha free transition normal pressure test}And zero lift angle of attack alpha_{0_ free transition _ Normal pressure test}；

Step A4: in the normal-pressure wind tunnel test state, a transition tape is pasted on different chord-direction positions of the surface of the wing to carry out forced transition to obtain the gradient C of the lift line_LαAnd zero lift angle of attack alpha₀A position change rule curve along with transition; then obtaining the corresponding lifting line slope C through the step A3_{L alpha free transition normal pressure test}And zero lift angle of attack alpha_{0_ free transition _ Normal pressure test}The slope C of the lift line obtained in this step_LαAnd zero lift angle of attack alpha₀Interpolating along with the change rule curve of transition position to obtain the free transition position Xn_{Free transition normal pressure test}The position can be regarded as the free transition position corresponding to the critical Re number;

step A5: free transition position Xn obtained by step A4_{Free transition normal pressure test}Adding the difference Δ Xn obtained in step A2 to obtain a new free transition position Xn_{'forced transition' normal pressure test}The position is the corresponding free transition position under the real flight altitude and the flight M number; by bringing a new free transition position Xn_{'forced transition' normal pressure test}Slope at the lifting line C_LαAnd zero lift angle of attack alpha₀Interpolation is carried out on the curve of the position change rule along with transition to obtain a corresponding new rising line slope C_{L alpha _ forced transition _ normal pressure test}And zero lift angle of attack alpha_{0_ forced transitionOrdinary pressure test}The slope C of the lifting line can be obtained_{L alpha _ forced transition _ normal pressure test}Slope C of the lift line_{L alpha free transition normal pressure test}Delta of (D) Δ C_{L alpha _ forced transition}Zero lift angle of attack alpha_{0_ forced transition _ Normal pressure test}Angle of attack alpha of zero lift_{0_ free transition _ Normal pressure test}Delta Δ α of (1)_{0_ forced transition}；

Step A6: from zero lift angle delta alpha_{0_ forced transition}Cruise angle of attack alpha and zero lift angle of attack alpha_{0_ free transition _ Normal pressure test}The difference Δ C of the slope of the lifting line_{L alpha _ forced transition}Slope of lift line C_{L alpha free transition normal pressure test}Calculating lift coefficient C_LCorrection amount of (1) (Δ C)_{L _ Compulsory transition}I.e. Δ C_{L _ Compulsory transition}＝-Δα_{0_ forced transition}×C_{L alpha free transition normal pressure test}+ΔC_{L alpha _ forced transition}×(α-α_{0_ free transition _ Normal pressure test})；

Step A7: superimposing the lift coefficient C of the reference state in step A1_{L _ boost}And coefficient of lift C in step A6_LCorrection amount of (1) (Δ C)_{L _ Compulsory transition}I.e. C_{L_Re}＝C_{L _ boost}+ΔC_{L _ Compulsory transition}(ii) a Obtaining the corrected lift coefficient C_{L_Re}And finishing the correction of the lift coefficient.

9. The method for modifying the Re effect of the flow transition-based supercritical laminar flow composite airfoil profile according to claim 7, wherein: the minimum coefficient of resistance C_DminThe specific steps of the correction are as follows:

step B1: increasing the proper pressure during the test to increase the Re number to the critical Re number or more and obtain the minimum resistance coefficient C of the reference state_{D _ supercharging}；

Step B2: determining the corrected flight height and flight M number, calculating the corresponding real Re number, and obtaining the free transition position Xn corresponding to the real Re number by combining the change curve of the laminar flow region of the upper surface of the wing along with the Re number in the real flight state_{Free transition (CFD)}Simultaneously obtaining the free transition position Xn corresponding to the critical Re number_{Free transition critical}To find the free transition position Xn_{Free transition (CFD)}And a free transition position Xn_{Free transition critical}The difference Δ Xn of (d);

step B3: carrying out normal-pressure wind tunnel test on the laminar flow wing type airplane to obtain the aerodynamic force characteristic of the airplane in the free transition state and obtain the corresponding minimum resistance coefficient C_{D alpha free transition normal pressure test}And zero lift angle of attack alpha_{0_ free transition _ Normal pressure test}；

Step B4: in the normal-pressure wind tunnel test state, the minimum resistance coefficient C is obtained by forcibly transition by sticking transition strips at different chord-direction positions on the surface of the wing_DαAnd zero lift angle of attack alpha₀A position change rule curve along with transition; then obtaining the corresponding minimum resistance coefficient C through the step B3_{D alpha free transition normal pressure test}And zero lift angle of attack alpha_{0_ free transition _ Normal pressure test}The minimum resistance coefficient C obtained in this step_DαAnd zero lift angle of attack alpha₀Interpolating along with the change rule curve of transition position to obtain the free transition position Xn_{Free transition normal pressure test}The position can be regarded as the free transition position corresponding to the critical Re number;

step B5: free transition position Xn obtained by step B4_{Free transition normal pressure test}Adding the difference Δ Xn obtained in step B2 to obtain a new free transition position Xn_{'forced transition' normal pressure test}The position is the corresponding free transition position under the real flight altitude and the flight M number; by bringing a new free transition position Xn_{'forced transition' normal pressure test}At the minimum drag coefficient C_DαAnd zero lift angle of attack alpha₀Interpolation is carried out on the curve of the position change rule along with transition to obtain a corresponding new minimum resistance coefficient C_{D alpha _ forced transition _ normal pressure test}And zero lift angle of attack alpha_{0_ forced transition _ Normal pressure test}The minimum resistance coefficient C can be obtained_{D alpha _ forced transition _ normal pressure test}And the minimum resistance coefficient C_{D alpha free transition normal pressure test}Delta of (D) Δ C_{D alpha _ forced transition}Zero lift angle of attack alpha_{0_ forced transition _ Normal pressure test}Angle of attack alpha of zero lift_{0_ free transition _ Normal pressure test}Delta Δ α of (1)_{0_ ForceTransition piece}；

Step B6: from zero lift angle delta alpha_{0_ forced transition}Cruise angle of attack alpha and zero lift angle of attack alpha_{0_ free transition _ Normal pressure test}The minimum resistance coefficient difference DeltaC_{D alpha _ forced transition}Minimum coefficient of resistance C_{D alpha free transition normal pressure test}Calculating the minimum resistance coefficient C_DCorrection amount of (1) (Δ C)_{D _ Compulsory transition}I.e. Δ C_{D _ Compulsory transition}＝-Δα_{0_ forced transition}×C_{D alpha free transition normal pressure test}+ΔC_{D alpha _ forced transition}×(α-α_{0_ free transition _ Normal pressure test})；

Step B7: superimposing the minimum drag coefficient C of the reference state in step B1_{D _ supercharging}And the minimum drag coefficient C in step B6_DCorrection amount of (1) (Δ C)_{D _ Compulsory transition}I.e. C_{D_Re}＝C_{D _ supercharging}+ΔC_{D _ Compulsory transition}(ii) a Obtaining the corrected minimum resistance coefficient C_{D_Re}And finishing the correction of the minimum resistance coefficient.

10. The method for modifying the Re effect of the flow transition-based supercritical laminar flow composite airfoil profile according to claim 7, wherein: the longitudinal moment coefficient C_mThe specific steps of the correction are as follows:

step C1: increasing proper pressure in the test process to increase the Re number to be tested to the critical Re number or more to the critical Re number, and acquiring the longitudinal moment coefficient C of the reference state_{m _ boost}；

Step C2: determining the corrected flight height and flight M number, calculating the corresponding real Re number, and obtaining the free transition position Xn corresponding to the real Re number by combining the change curve of the laminar flow region of the upper surface of the wing along with the Re number in the real flight state_{Free transition (CFD)}Simultaneously obtaining the free transition position Xn corresponding to the critical Re number_{Free transition critical}To find the free transition position Xn_{Free transition (CFD)}And a free transition position Xn_{Free transition critical}The difference Δ Xn of (d);

step C3: carrying out normal-pressure wind tunnel test on the laminar flow wing type airplane to obtain the aerodynamics of the airplane in the free transition stateForce characteristics to obtain corresponding longitudinal moment coefficient C_{m alpha free transition normal pressure test}And zero lift moment coefficient C_{m (alpha is 0) _ free transition _ normal pressure test}；

Step C4: in the normal-pressure wind tunnel test state, the longitudinal moment coefficient C is obtained by forcibly transition by sticking transition strips at different chord-direction positions on the surface of the wing_mαAnd zero lift moment coefficient C_m(α＝0)A position change rule curve along with transition; then the corresponding longitudinal moment coefficient C obtained in the step C3 is used_{m alpha free transition normal pressure test}And zero lift moment coefficient C_{m (alpha is 0) _ free transition _ normal pressure test}The longitudinal moment coefficient C obtained in this step_mαInterpolating along with the change rule curve of transition position to obtain the free transition position Xn_{Free transition normal pressure test}The position can be regarded as the free transition position corresponding to the critical Re number;

step C5: the free transition position Xn obtained in step C4_{Free transition normal pressure test}Adding the difference Δ Xn obtained in step C2 to obtain a new free transition position Xn_{'forced transition' normal pressure test}The position is the corresponding free transition position under the real flight altitude and the flight M number; by bringing a new free transition position Xn_{'forced transition' normal pressure test}In the longitudinal direction of the moment coefficient C_mαInterpolation is carried out on the curve of the position change rule along with transition to obtain a corresponding new longitudinal moment coefficient C_{m alpha _ forced transition _ normal pressure test}And a new zero lift moment C_{m (alpha is 0) _ forced transition _ normal pressure test}The longitudinal moment coefficient C can be obtained_{m alpha _ forced transition _ normal pressure test}Coefficient of longitudinal moment C_{m alpha free transition normal pressure test}Delta of (D) Δ C_{m alpha _ forced transition}(ii) a While obtaining zero lift moment coefficient C_{m (alpha is 0) _ forced transition _ normal pressure test}Coefficient of zero lift moment C_{m (alpha is 0) _ free transition _ normal pressure test}Delta of (D) Δ C_{m (alpha 0) _ forced transition}；

Step C6: from cruise angle of attack alpha, zero lift moment coefficient difference delta C_{m (alpha 0) _ forced transition}Longitudinal moment coefficient difference Delta C_{m alpha _ forced transition}Calculating the longitudinal moment coefficient C_mCorrection amount of (1) (Δ C)_{m _ Compulsory transition}I.e. Δ C_{m _ Compulsory transition}＝ΔC_{m (alpha 0) _ forced transition}+ΔC_{m alpha _ forced transition}×α；

Step C7: superimposing the longitudinal moment coefficient C of the reference state in step C1_{m _ boost}And the longitudinal moment coefficient C in step C6_mCorrection amount of (1) (Δ C)_{m _ Compulsory transition}I.e. C_{m_Re}＝C_{m _ boost}+ΔC_{m _ Compulsory transition}(ii) a Obtaining the corrected longitudinal moment coefficient C_{m_Re}And finishing the correction of the longitudinal moment coefficient.

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